It is generally known in the art to use stringer-stiffened structures made of fiber composite materials in the field of aircraft construction. For example, these may be used as primary structures of a passenger aircraft, whereby the primary application known in the art is the empennage or rudder and vertical stabilizer assembly of known Airbus aircraft, which is completely fabricated of carbon fiber reinforced composite materials. In this context, two publications each provide certain impressions or ideas for realizing the known stringer-stiffened structures using fiber composite materials. See U. Bieling, "Serieneinsatz von Faserverbundwerkstoffen im Flugzeugbau--dargestellt am Seitenleitwerk des Airbus", "Series Application of Fiber Composite Materials in Aircraft Construction--Represented in the Rudder and Vertical Stabilizer Assembly of the Airbus", VDI Berichte No. 965.1, pages 77 to 88, VDI publishers Dusseldorf, Germany, 1992; and also J. Rouchon, "Certification of Large Aircraft Composite Structures, Recent Progress and New Trends in Compliance Philosophy", Proceedings of the 17th ICAS Conference Stockholm, Sweden, 1990.
In part, Bieling discusses the fabrication of stiffening internal structures of highly integrated carbon fiber composite components for a middle shell or middle box of a rudder and stabilizer assembly. Among other things, Bieling also addresses the complex fabrication that is required in the so-called modular core technology. In this context, Bieling discusses in detail the fabrication process for a middle box shell of a rudder and stabilizer assembly, in which the carbon fiber composite pre-cut blanks that are supplied in a sorted manner must first be separated into skin layers and so-called bandages. The skin layers are manually laid into laminating molds, while the bandages are wrapped or wound around the so-called modular cores in a partially automated process. Multiple layers of these bandages are wrapped around the cores in a single process operation, when rectangular modular cores are used for the shells of the rudder and stabilizer assembly. The entire wrapping or winding process involves high technical demands, especially requiring a sufficiently high and reproducible adhesive bonding characteristic of the prepregs. In this context, trapezoidal or geometrically complex modular cores must be manually wrapped layer by layer.
Thereafter, the wrapped modular cores are arranged in a row in an exactly prescribed order on a rotatable support. In this context, the prepregs arranged between the lined-up modular cores, after curing, form the stiffening internal structure of the highly integrated carbon fiber composite components. After completion of a subsequent laminating process for the skin layers and fixing of all the modular cores, the rotatable support is rotated in such a manner that all of the cores hang downward, and then the rotatable support is lowered down onto the skin layers laying in the laminating mold. After releasing the cores from the rotatable support, the cores remain on the laminating form and the component is made ready for the curing process.
After the curing has been completed, the components are removed from the forming mold. In this context, first the middle parts of the three-part modular cores and thereafter the side parts are removed or pulled out of the component. After all of the core parts have been removed, the component is lifted out of the forming mold and subjected further to a mechanical machining. The modular cores are cleaned of any residual resin and then once again provided with separating members. The laminating mold is cleaned and provided with separating members. To repeat the above described steps of the process, the individual core parts are then joined together again to form a modular core which is used for carrying out the next wrapping process. After all of the above mentioned steps have been carried out, the components are subsequently mechanically machined, for example by boring, milling, perimeter machining or flash removal or the like, and are then subjected to a non-destructive testing.
It is desirable to use fiber composite materials for the primary structures such as an empennage, e.g. a rudder and stabilizer assembly, of an aircraft in view of the many advantages such as a weight reduction, very good specific strength and stiffness characteristics, a reduced total number of individual components, very good corrosion resistance, very good fatigue performance, among others. However, from the above description of the known fabrication processes, it is clear that further improvements in the fabrication of such primary structures are desirable or even necessary to achieve a rational fabrication with a low expense and effort in the way of tooling, equipment, and process steps, while reliably meeting the high technical demands that are typical in the field of aircraft construction. These considerations especially apply in the construction of high capacity civilian commercial aircraft. A further subordinate requirement is the use of rationalized production technologies for achieving stiffened primary structures using fiber composite materials, which may be applied to empennage structures as well as the wings of aircraft.
For the above reasons, it is important in the field to effectively accommodate or adapt known production techniques in order to achieve the simplest possible application of stiffening members such as stringers onto complexly curved shells or skin layers. In this context, it is generally known in the field that significant problems arise when fitting stiffening members such as stringers onto the curved shells or skin layers in the production of stringer-stiffened structures of fiber composite materials. When pre-fabricated and pre-cured stiffening members are combined with a shell or skin structure having a relatively sharp curvature, various assembly and fabrication difficulties arise in achieving a proper alignment and conformance of the parts, and ultimately a warping deformation of the finished component results due to imperfect matching or conformance of the various parts. On the other hand, any attempt to use non-cured semifinished parts for the stiffening members have previously always required the use of very complicated and costly molds or forming tools for properly molding the stiffening members, bonding or joining the stiffening members onto the shell structure, and curing the stiffening members. The above mentioned publications of Bieling and Rouchon provide no disclosure or suggestions toward how the fabrication of complexly shaped stringer-stiffened structures can be improved while simultaneously avoiding the above mentioned problems.